US20130205792A1 - Cooling hole with asymmetric diffuser - Google Patents
Cooling hole with asymmetric diffuser Download PDFInfo
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- US20130205792A1 US20130205792A1 US13/544,136 US201213544136A US2013205792A1 US 20130205792 A1 US20130205792 A1 US 20130205792A1 US 201213544136 A US201213544136 A US 201213544136A US 2013205792 A1 US2013205792 A1 US 2013205792A1
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- lobe
- wall
- metering section
- section
- component
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
- F01D5/12—Blades
- F01D5/14—Form or construction
- F01D5/18—Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
- F01D5/186—Film cooling
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D9/00—Stators
- F01D9/06—Fluid supply conduits to nozzles or the like
- F01D9/065—Fluid supply or removal conduits traversing the working fluid flow, e.g. for lubrication-, cooling-, or sealing fluids
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
- F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
- F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
- F23R3/002—Wall structures
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
- F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
- F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
- F23R3/02—Continuous combustion chambers using liquid or gaseous fuel characterised by the air-flow or gas-flow configuration
- F23R3/04—Air inlet arrangements
- F23R3/06—Arrangement of apertures along the flame tube
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2240/00—Components
- F05D2240/80—Platforms for stationary or moving blades
- F05D2240/81—Cooled platforms
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2250/00—Geometry
- F05D2250/30—Arrangement of components
- F05D2250/32—Arrangement of components according to their shape
- F05D2250/324—Arrangement of components according to their shape divergent
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2250/00—Geometry
- F05D2250/70—Shape
- F05D2250/73—Shape asymmetric
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2260/00—Function
- F05D2260/20—Heat transfer, e.g. cooling
- F05D2260/202—Heat transfer, e.g. cooling by film cooling
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
- F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
- F23R2900/00—Special features of, or arrangements for continuous combustion chambers; Combustion processes therefor
- F23R2900/00018—Manufacturing combustion chamber liners or subparts
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
- F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
- F23R2900/00—Special features of, or arrangements for continuous combustion chambers; Combustion processes therefor
- F23R2900/03042—Film cooled combustion chamber walls or domes
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- Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
- Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
- Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
- Y02T50/00—Aeronautics or air transport
- Y02T50/60—Efficient propulsion technologies, e.g. for aircraft
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- Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
- Y10—TECHNICAL SUBJECTS COVERED BY FORMER USPC
- Y10T—TECHNICAL SUBJECTS COVERED BY FORMER US CLASSIFICATION
- Y10T29/00—Metal working
- Y10T29/49—Method of mechanical manufacture
- Y10T29/49316—Impeller making
- Y10T29/4932—Turbomachine making
- Y10T29/49323—Assembling fluid flow directing devices, e.g., stators, diaphragms, nozzles
Definitions
- This invention relates generally to turbomachinery, and specifically to turbine flow path components for gas turbine engines.
- the invention relates to cooling techniques for airfoils and other gas turbine engine components exposed to hot working fluid flow, including, but not limited to, rotor blades and stator vane airfoils, endwall surfaces including platforms, shrouds and compressor and turbine casings, combustor liners, turbine exhaust assemblies, thrust augmentors and exhaust nozzles.
- Gas turbine engines are rotary-type combustion turbine engines built around a power core made up of a compressor, combustor and turbine, arranged in flow series with an upstream inlet and downstream exhaust.
- the compressor section compresses air from the inlet, which is mixed with fuel in the combustor and ignited to generate hot combustion gas.
- the turbine section extracts energy from the expanding combustion gas, and drives the compressor section via a common shaft. Expanded combustion products are exhausted downstream, and energy is delivered in the form of rotational energy in the shaft, reactive thrust from the exhaust, or both.
- Gas turbine engines provide efficient, reliable power for a wide range of applications in aviation, transportation and industrial power generation.
- Small-scale gas turbine engines typically utilize a one-spool design, with co-rotating compressor and turbine sections.
- Larger-scale combustion turbines including jet engines and industrial gas turbines (IGTs) are generally arranged into a number of coaxially nested spools. The spools operate at different pressures, temperatures and spool speeds, and may rotate in different directions.
- Individual compressor and turbine sections in each spool may also be subdivided into a number of stages, formed of alternating rows of rotor blade and stator vane airfoils.
- the airfoils are shaped to turn, accelerate and compress the working fluid flow, or to generate lift for conversion to rotational energy in the turbine.
- Industrial gas turbines often utilize complex nested spool configurations, and deliver power via an output shaft coupled to an electrical generator or other load, typically using an external gearbox.
- CCGTs combined cycle gas turbines
- a steam turbine or other secondary system is used to extract additional energy from the exhaust, improving thermodynamic efficiency.
- Gas turbine engines are also used in marine and land-based applications, including naval vessels, trains and armored vehicles, and in smaller-scale applications such as auxiliary power units.
- turbojet engines thrust is generated primarily from the exhaust.
- Modern fixed-wing aircraft generally employ turbofan and turboprop configurations, in which the low pressure spool is coupled to a propulsion fan or propeller.
- Turboshaft engines are employed on rotary-wing aircraft, including helicopters, typically using a reduction gearbox to control blade speed.
- Unducted (open rotor) turbofans and ducted propeller engines also known, in a variety of single-rotor and contra-rotating designs with both forward and aft mounting configurations.
- Aviation turbines generally utilize two and three-spool configurations, with a corresponding number of coaxially rotating turbine and compressor sections.
- the high pressure turbine drives a high pressure compressor, forming the high pressure spool or high spool.
- the low-pressure turbine drives the low spool and fan section, or a shaft for a rotor or propeller.
- three-spool engines there is also an intermediate pressure spool.
- Aviation turbines are also used to power auxiliary devices including electrical generators, hydraulic pumps and elements of the environmental control system, for example using bleed air from the compressor or via an accessory gearbox.
- Additional turbine engine applications and turbine engine types include intercooled, regenerated or recuperated and variable cycle gas turbine engines, and combinations thereof.
- these applications include intercooled turbine engines, for example with a relatively higher pressure ratio, regenerated or recuperated gas turbine engines, for example with a relatively lower pressure ratio or for smaller-scale applications, and variable cycle gas turbine engines, for example for operation under a range of flight conditions including subsonic, transonic and supersonic speeds.
- Combined intercooled and regenerated/recuperated engines are also known, in a variety of spool configurations with traditional and variable cycle modes of operation.
- Turbofan engines are commonly divided into high and low bypass configurations.
- High bypass turbofans generate thrust primarily from the fan, which accelerates airflow through a bypass duct oriented around the engine core. This design is common on commercial aircraft and transports, where noise and fuel efficiency are primary concerns.
- the fan rotor may also operate as a first stage compressor, or as a pre-compressor stage for the low-pressure compressor or booster module.
- Variable-area nozzle surfaces can also be deployed to regulate the bypass pressure and improve fan performance, for example during takeoff and landing.
- Advanced turbofan engines may also utilize a geared fan drive mechanism to provide greater speed control, reducing noise and increasing engine efficiency, or to increase or decrease specific thrust.
- Low bypass turbofans produce proportionally more thrust from the exhaust flow, generating greater specific thrust for use in high-performance applications including supersonic jet aircraft.
- Low bypass turbofan engines may also include variable-area exhaust nozzles and afterburner or augmentor assemblies for flow regulation and short-term thrust enhancement.
- Specialized high-speed applications include continuously afterburning engines and hybrid turbojet/ramjet configurations.
- turbine performance depends on the balance between higher pressure ratios and core gas path temperatures, which tend to increase efficiency, and the related effects on service life and reliability due to increased stress and wear. This balance is particularly relevant to gas turbine engine components in the hot sections of the compressor, combustor, turbine and exhaust sections, where active cooling is required to prevent damage due to high gas path temperatures and pressures.
- Components present in the hot gas path of a gas turbine engine require cooling to prevent component melting and to reduce the effects of thermal fatigue and wear.
- Hollow blades and vanes, combustor walls and other components include thin metal walls made of high strength materials that provide durability. While these materials reduce the amount of cooling necessary, components in the hot gas path still require some sort of surface cooling.
- Film cooling holes are often used to cool these components. This type of cooling works by delivering cool air (e.g., air bled from a compressor) through small holes in the wall surface of the component. This air creates a thin layer (film) of cool air on the surface of the component wall, protecting it from higher temperature air and gases.
- cool air e.g., air bled from a compressor
- film thin layer of cool air on the surface of the component wall, protecting it from higher temperature air and gases.
- One consideration with film cooling is that injecting cool air into a component reduces engine efficiency. The drop in efficiency increases as the amount of cooling airflow increases.
- Diffusion cooling holes were designed to increase the spread of the cooling film to reduce the debit on engine efficiency. By spreading out the film of cooling air, smaller amounts of cooling air could be used to cool an area.
- One problem with diffusion cooling holes is flow separation. Diffusion cooling holes can only spread cooling air to a certain extent before the flow separates, creating a “hole” in the cooling film. Flow separation is likely to occur at the “corners” of state of the art diffusion holes. Additionally, at high blowing ratios, the cooling film can “jet” or “blow off” the surface of the component, allowing nearby hot gases to cover the surface and reducing cooling effectiveness.
- a gas turbine engine component includes a wall having first and second wall surfaces and a cooling hole extending through the wall.
- the cooling hole includes an inlet located at the first wall surface, an outlet located at the second wall surface, a metering section extending downstream from the inlet and a diffusing section extending from the metering section to the outlet.
- the diffusing section includes a first lobe diverging longitudinally from the metering section and a second lobe adjacent the first lobe and diverging longitudinally and laterally from the metering section.
- a wall of a component of a gas turbine engine includes first and second wall surfaces, an inlet located at the first wall surface, an outlet located at the second wall surface, a metering section commencing at the inlet and extending downstream from the inlet and a diffusing section extending from the metering section and terminating at the outlet.
- the diffusing section includes a first lobe diverging longitudinally from the metering section, a second lobe adjacent the first lobe and diverging longitudinally and laterally from the metering section and a ridge located between the first and second lobes.
- a method for producing a cooling hole in a gas turbine engine wall having first and second wall surfaces includes forming a metering section between the first wall surface and the second wall surface and forming a diffusing section between the metering section and the second wall surface.
- the diffusing section includes a first lobe in line with the metering section and a second lobe that diverges laterally from the metering section.
- the diffusing section distributes the flow of the fluid into the lobes to form a film of cooling fluid at a hole outlet at the second wall surface of the gas turbine engine wall.
- FIG. 1 is a cross-sectional view of a gas turbine engine.
- FIG. 2A is a perspective view of an airfoil for the gas turbine engine, in a rotor blade configuration.
- FIG. 2B is a perspective view of an airfoil for the gas turbine engine, in a stator vane configuration.
- FIG. 3 is a view of a wall having cooling holes with asymmetric diffusing sections.
- FIG. 4 is a sectional view of the cooling hole of FIG. 3 taken along the line 4 - 4 .
- FIG. 5 is a view of the cooling hole of FIG. 4 taken along the line 5 - 5 .
- FIG. 5A is a view of the cooling hole of FIG. 5 taken along the line A-A.
- FIG. 6 is another embodiment of a cooling hole with an asymmetric diffusing section.
- FIG. 7 is a sectional view of another embodiment of a cooling hole with an asymmetric diffusing section.
- FIG. 8 is a view of the cooling hole of FIG. 7 taken along the line 8 - 8 .
- FIG. 9A is a simplified flow diagram illustrating one embodiment of a method for producing a tri-lobed cooling hole.
- FIG. 9B is a simplified flow diagram illustrating another embodiment of a method for producing a tri-lobed cooling hole.
- FIG. 1 is a cross-sectional view of gas turbine engine 10 .
- Gas turbine engine (or turbine engine) 10 includes a power core with compressor section 12 , combustor 14 and turbine section 16 arranged in flow series between upstream inlet 18 and downstream exhaust 20 .
- Compressor section 12 and turbine section 16 are arranged into a number of alternating stages of rotor airfoils (or blades) 22 and stator airfoils (or vanes) 24 .
- propulsion fan 26 is positioned in bypass duct 28 , which is coaxially oriented about the engine core along centerline (or turbine axis) C L .
- An open-rotor propulsion stage 26 may also provided, with turbine engine 10 operating as a turboprop or unducted turbofan engine.
- fan rotor 26 and bypass duct 28 may be absent, with turbine engine 10 configured as a turbojet or turboshaft engine, or an industrial gas turbine.
- components of gas turbine engine 10 are provided with an improved cooling configuration, as described below.
- Suitable components for the cooling configuration include rotor airfoils 22 , stator airfoils 24 and other gas turbine engine components exposed to hot gas flow, including, but not limited to, platforms, shrouds, casings and other endwall surfaces in hot sections of compressor 12 and turbine 16 , and liners, nozzles, afterburners, augmentors and other gas wall components in combustor 14 and exhaust section 20 .
- compressor section 12 includes low pressure compressor (LPC) 30 and high pressure compressor (HPC) 32
- turbine section 16 includes high pressure turbine (HPT) 34 and low pressure turbine (LPT) 36
- Low pressure compressor 30 is rotationally coupled to low pressure turbine 36 via low pressure (LP) shaft 38 , forming the LP spool or low spool.
- High pressure compressor 32 is rotationally coupled to high pressure turbine 34 via high pressure (HP) shaft 40 , forming the HP spool or high spool.
- Flow F at inlet 18 divides into primary (core) flow F P and secondary (bypass) flow F S downstream of fan rotor 26 .
- Fan rotor 26 accelerates secondary flow F S through bypass duct 28 , with fan exit guide vanes (FEGVs) 42 to reduce swirl and improve thrust performance.
- FEGVs fan exit guide vanes
- structural guide vanes (SGVs) 42 are used, providing combined flow turning and load bearing capabilities.
- Primary flow F P is compressed in low pressure compressor 30 and high pressure compressor 32 , then mixed with fuel in combustor 14 and ignited to generate hot combustion gas.
- the combustion gas expands to provide rotational energy in high pressure turbine 34 and low pressure turbine 36 , driving high pressure compressor 32 and low pressure compressor 30 , respectively.
- Expanded combustion gases exit through exhaust section (or exhaust nozzle) 20 , which can be shaped or actuated to regulate the exhaust flow and improve thrust performance.
- Low pressure shaft 38 and high pressure shaft 40 are mounted coaxially about centerline C L , and rotate at different speeds.
- Fan rotor (or other propulsion stage) 26 is rotationally coupled to low pressure shaft 38 .
- fan drive gear system 44 is provided for additional fan speed control, improving thrust performance and efficiency with reduced noise output.
- Fan rotor 26 may also function as a first-stage compressor for gas turbine engine 10 , and LPC 30 may be configured as an intermediate compressor or booster.
- propulsion stage 26 has an open rotor design, or is absent, as described above.
- Gas turbine engine 10 thus encompasses a wide range of different shaft, spool and turbine engine configurations, including one, two and three-spool turboprop and (high or low bypass) turbofan engines, turboshaft engines, turbojet engines, and multi-spool industrial gas turbines.
- turbine efficiency and performance depend on the overall pressure ratio, defined by the total pressure at inlet 18 as compared to the exit pressure of compressor section 12 , for example at the outlet of high pressure compressor 32 , entering combustor 14 .
- Higher pressure ratios also result in greater gas path temperatures, increasing the cooling loads on rotor airfoils 22 , stator airfoils 24 and other components of gas turbine engine 10 .
- Suitable components include, but are not limited to, cooled gas turbine engine components in compressor sections 30 and 32 , combustor 14 , turbine sections 34 and 36 , and exhaust section 20 of gas turbine engine 10 .
- FIG. 2A is a perspective view of rotor airfoil (or blade) 22 for gas turbine engine 10 , as shown in FIG. 1 , or for another turbomachine.
- Rotor airfoil 22 extends axially from leading edge 51 to trailing edge 52 , defining pressure surface 53 (front) and suction surface 54 (back) therebetween.
- Pressure and suction surfaces 53 and 54 form the major opposing surfaces or walls of airfoil 22 , extending axially between leading edge 51 and trailing edge 52 , and radially from root section 55 , adjacent inner diameter (ID) platform 56 , to tip section 57 , opposite ID platform 56 .
- tip section 57 is shrouded.
- Cooling holes or outlets 60 are provided on one or more surfaces of airfoil 22 , for example along leading edge 51 , trailing edge 52 , pressure (or concave) surface 53 , or suction (or convex) surface 54 , or a combination thereof. Cooling holes or passages 60 may also be provided on the endwall surfaces of airfoil 22 , for example along ID platform 56 , or on a shroud or engine casing adjacent tip section 57 .
- FIG. 2B is a perspective view of stator airfoil (or vane) 24 for gas turbine engine 10 , as shown in FIG. 1 , or for another turbomachine.
- Stator airfoil 24 extends axially from leading edge 61 to trailing edge 62 , defining pressure surface 63 (front) and suction surface 64 (back) therebetween.
- Pressure and suction surfaces 63 and 64 extend from inner (or root) section 65 , adjacent ID platform 66 , to outer (or tip) section 67 , adjacent outer diameter (OD) platform 68 .
- Cooling holes or outlets 60 are provided along one or more surfaces of airfoil 24 , for example leading or trailing edge 61 or 62 , pressure (concave) or suction (convex) surface 63 or 64 , or a combination thereof. Cooling holes or passages 60 may also be provided on the endwall surfaces of airfoil 24 , for example along ID platform 66 and OD platform 68 .
- Rotor airfoils 22 ( FIG. 2A ) and stator airfoils 24 ( FIG. 2B ) are formed of high strength, heat resistant materials such as high temperature alloys and superalloys, and are provided with thermal and erosion-resistant coatings.
- Airfoils 22 and 24 are also provided with internal cooling passages and cooling holes 60 to reduce thermal fatigue and wear, and to prevent melting when exposed to hot gas flow in the higher temperature regions of a gas turbine engine or other turbomachine.
- Cooling holes 60 deliver cooling fluid (e.g., steam or air from a compressor) through the outer walls and platform structures of airfoils 22 and 24 , creating a thin layer (or film) of cooling fluid to protect the outer (gas path) surfaces from high temperature flow.
- cooling fluid e.g., steam or air from a compressor
- Cooling holes 60 are thus provided with improved metering and inlet geometry to reduce jets and blow off, and improved diffusion and exit geometry to reduce flow separation and corner effects. Cooling holes 60 reduce flow requirements and improve the spread of cooling fluid across the hot outer surfaces of airfoils 22 and 24 , and other gas turbine engine components, so that less flow is needed for cooling and efficiency is maintained or increased.
- FIG. 3 illustrates a view of a wall of a gas turbine engine component having cooling holes with asymmetric diffusing sections.
- Wall 100 includes first wall surface 102 and second wall surface 104 .
- wall 100 is primarily metallic and second wall surface 104 can include a thermal barrier coating.
- Cooling holes 106 are oriented so that their inlets are positioned on the first wall surface 102 and their outlets are positioned on second wall surface 104 .
- second wall surface 104 is in proximity to high temperature gases (e.g., combustion gases, hot air). Cooling air is delivered inside wall 100 where it exits the interior of the component through cooling holes 106 and forms a cooling film on second wall surface 104 .
- cooling holes 106 have two lobes in the diffusing section of the cooling hole outlet on second wall surface 104 .
- cooling air flows out of cooling holes 106 and flows through each of the lobes in the diffusing section.
- Cooling holes 106 can be arranged in a row on wall 100 as shown in FIG. 3 and positioned axially so that the cooling air flows in substantially the same direction longitudinally as the high temperature gases flowing past wall 100 .
- cooling air passing through cooling holes 106 exits cooling holes traveling in substantially the same direction as the high temperature gases flowing along second wall surface 104 (represented by arrow H).
- the linear row of cooling holes 106 is substantially perpendicular to the direction of flow H.
- the orientation of cooling holes 106 can be arranged on second wall surface 104 so that the flow of cooling air is at an angle between parallel and perpendicular.
- Cooling holes 106 can also be provided in a staggered formation on wall 100 . Cooling holes 106 can be located on a variety of components that require cooling. Suitable components include, but are not limited to, turbine vanes and blades, combustors, blade outer air seals, augmentors, etc. Cooling holes 106 can be located on the pressure side or suction side of vanes and blades. Cooling holes 106 can also be located on the blade tip or blade or vane platforms. Cooling holes 106 can also be located near airfoil endwalls or at other locations and individually aligned to provide targeted flow of cooling air.
- FIGS. 4 and 5 illustrate embodiments of cooling hole 106 in greater detail.
- FIG. 4 illustrates a sectional view of cooling hole 106 of FIG. 3 taken along the line 4 - 4 .
- FIG. 5 illustrates a view of cooling hole 106 of FIG. 4 taken along the line 5 - 5 .
- Cooling hole 106 includes inlet 110 , metering section 112 and diffusing section 114 .
- Inlet 110 is an opening located on first wall surface 102 .
- Cooling air C enters cooling hole 106 through inlet 110 and passes through metering section 112 and diffusing section 114 before exiting cooling hole 106 at outlet 116 along second wall surface 104 .
- Metering section 112 extends downstream from inlet 110 and controls (meters) the flow of cooling air through cooling hole 106 .
- metering section 112 has a substantially constant flow area from inlet 110 to diffusing section 114 .
- Metering section 112 can have circular, oblong (oval or elliptical) or racetrack (oval with two parallel sides having straight portions) shaped cross sections. In FIGS. 4 and 5 , metering section 112 has a circular cross section.
- Circular metering sections 112 have a length l and diameter d. In some embodiments, circular metering section 112 has a length l according to the relationship: d ⁇ l ⁇ 3d.
- the length of metering section 112 is between one and three times its diameter.
- the length of metering section 22 can exceed 3d, reaching upwards of 30d.
- metering section 112 has an oblong or racetrack-shaped or other shaped cross section. As oblong and racetrack configurations are not circular, their metering sections 112 have a length l and hydraulic diameter d h . In some embodiments, metering section 112 has a length l according to the relationship: d h ⁇ l ⁇ 3d h . That is, the length of metering section 112 is between one and three times its hydraulic diameter.
- metering section 112 can exceed 3d h , reaching upwards of 30d h .
- metering section 112 is inclined with respect to wall 100 as illustrated in FIG. 4 (i.e. metering section 112 is not perpendicular to wall 100 ).
- Metering section 112 has a longitudinal axis represented by numeral 118 .
- Metering section 112 also has a lateral sidewall 113 as shown in FIG. 5 .
- Diffusing section 114 is adjacent to and downstream from metering section 112 . Cooling air C diffuses within diffusing section 114 before exiting cooling hole 106 at outlet 116 along second wall surface 104 . Once cooling air C exits metering section 112 , the flow of air expands to fill diffusing section 114 . Cooling air C diffuses longitudinally (shown best in FIG. 4 ). In some embodiments, cooling air diffuses both longitudinally and laterally (shown best in FIG. 5 ) in diffusing section 114 . Second wall surface 104 includes upstream end 120 (upstream of cooling hole 106 ) and downstream end 122 (downstream from cooling hole 106 ). Diffusing section 114 opens along second wall surface 104 between upstream end 120 and downstream end 122 . As shown in FIG. 4 , cooling air C diffuses in diffusing section 114 as it flows towards outlet 116 .
- diffusing section 114 includes two lobes 124 and 126 .
- Each lobe 124 , 126 has a bottom surface (bottom surfaces 130 and 132 , respectively).
- Lobes 124 and 126 each have a side wall along the outer edge of diffusing section 114 (side walls 136 and 138 , respectively).
- Each lobe 124 , 126 also has a trailing edge (trailing edges 140 and 142 , respectively).
- Lobes 124 and 126 meet along ridge 146 .
- Ridge 146 can be straight or curved, both longitudinally and laterally. As shown in FIG. 4 , each lobe diverges longitudinally from metering section 112 .
- FIG. 4 each lobe diverges longitudinally from metering section 112 .
- FIG. 4 illustrates a sectional view taken through the center of cooling hole 106 and shows ridge 146 between lobes 124 and 126 .
- Ridge 146 is inclined with respect to second wall surface 104 as shown by inclination angle ⁇ 1 .
- Bottom surfaces 130 and 132 of lobes 124 and 126 are also inclined with respect to second wall surface 104 as shown by inclination angle ⁇ 2 .
- Inclination angle ⁇ 2 indicates a downstream angle for each lobe.
- bottom surfaces 130 and 132 of lobes 124 and 126 have the same inclination angle ⁇ 2 (downstream angle). As described in greater detail below, bottom surfaces 130 and 132 do not need to have the same depth or inclination angle.
- Cooling air C flowing through diffusing section 114 diverges longitudinally from longitudinal axis 118 as it “attaches” to bottom surfaces 130 and 132 of respective lobes 124 and 126 .
- Lobes 124 and 126 meet with second wall surface 104 at trailing edges 140 and 142 , respectively.
- cooling air C passing through cooling hole 106 also diffuses longitudinally near upstream end 120 .
- the upstream portion of diffusing section 114 is bounded by forward edge 150 .
- Forward edge 150 can be parallel with the upstream edge of metering section 112 (and with longitudinal axis 118 ), inclined towards upstream end 120 or inclined towards downstream end 122 .
- forward edge 150 is parallel with the upstream edge of metering section 112 (i.e. no upstream longitudinal diffusion) or inclined towards downstream end 122 .
- forward edge 150 is inclined slightly towards upstream end 120 from longitudinal axis 118 (represented by inclination angle ⁇ 3 ).
- forward edge 150 is inclined towards upstream end 120 to accommodate certain manufacturing methods. In these embodiments, the magnitude of inclination angle ⁇ 3 is minimized to less than about 15° and, in another embodiment, to less than about 1°. By minimizing inclination angle ⁇ 3 and positioning the end of forward edge 150 at second wall surface 104 as far downstream as possible, cooling air C exiting outlet 116 is likely to be more effective. In some embodiments, forward edge 150 is inclined towards downstream 122 (rather than upstream end 120 ) at an inclination angle ⁇ 3 of up to about ⁇ 2°.
- cooling air C diffuses longitudinally within diffusing section 114 as shown in FIG. 4
- cooling air also diffuses laterally within diffusing section 114 as shown in FIG. 5 .
- Lobe 126 diverges laterally with respect to metering section 112 .
- Lobe 124 includes side wall 136 on the side of lobe 124 opposite ridge 146 .
- Lobe 126 includes side wall 138 on the side of lobe 126 opposite ridge 148 .
- lobe 126 diverges laterally in a downward direction away from centerline axis 152 .
- Centerline axis 152 is a longitudinal axis passing through the center of metering section 112 .
- sidewall 136 of lobe 124 is parallel with lateral sidewall 113 of metering section 112 and does not diverge in an upward direction away from centerline axis 152 .
- lobe 124 does not laterally diverge away from centerline axis 152 to a substantial degree.
- Ridge 146 is angled downward (as shown in FIG. 5 ) slightly, allowing some lateral divergence of flow through lobe 124 .
- Ridge 146 can be angled downward to a greater degree to increase the lateral divergence of flow through lobe 124 in one direction.
- ridge 146 can be parallel to sidewall 136 and metering section 112 .
- Ridge 146 aids in directing cooling air C into lobes 124 and 126 .
- Ridge 146 is generally an inverted V-shaped portion where the adjacent lobes meet. Ridge 146 can form a sharp edge between the lobes, where edges of adjacent lobes meet at a point. Alternatively, ridge 146 can be rounded or have other geometric shapes. Ridge 146 can form a straight line between adjacent lobes. Alternatively, ridge 146 can be laterally curved. As cooling air C exits metering section 112 and enters diffusing section 114 , cooling air 26 encounters ridge 146 . Ridge 146 can extend farther towards second wall surface 104 than lobes 124 and 126 as shown in FIG.
- ridge 146 projects towards second wall surface 104 and serve to guide the flow of cooling air C into lobes 124 and 126 .
- Ridge 146 divides the flow of cooling air C between lobes 124 and 126 , causing cooling air C flowing into lobe 126 to diverge laterally to correspond to the shape of lobe 126 .
- bottom surfaces 130 and 132 of lobes 124 and 126 respectively, include a curved portion. As shown in FIG. 5A , the outer portion of lobes 124 and 126 can be curved.
- Lobe 124 includes a curved surface at side wall 136 and a curved bottom surface 130 .
- Lobe 126 includes a curved surface at side wall 138 and a curved bottom surface 132 .
- bottom surfaces 130 and 132 are concave (i.e. curve towards first wall surface 102 ).
- FIG. 6 is a top view of another embodiment of a cooling hole, cooling hole 106 A.
- diffusing section 114 of cooling hole 106 A includes three lobes.
- Lobe 128 is located between lobes 124 and 126 .
- Lobe 128 includes bottom surface 134 and trailing edge 144 .
- Ridge 147 separates lobe 124 and lobe 128
- ridge 148 separates lobe 126 and lobe 128 .
- Adding lobe 128 increases the amount of lateral divergence of cooling air C in diffusing section 114 .
- Ridges 147 and 148 divide the flow of cooling air C between lobes 124 , 126 and 128 , causing cooling air C flowing into lobes 126 and 128 to diverge laterally.
- Lobes 124 , 126 and 128 meet and blend with second wall surface 104 at trailing edges 140 , 142 and 144 , respectively.
- Lobes 124 , 126 and 128 can blend with second wall surface 104 in a number of ways.
- each lobe blends with second wall surface 104 at the same axial distance from inlet 110 , such that trailing edges 140 , 142 and 144 form a generally straight line.
- FIG. 5 illustrates an embodiment in which the trailing edges of the lobes form a generally straight line.
- trailing edges 140 , 142 and 144 are equidistant from a point on upstream end 120 .
- lobes 124 , 126 and 128 have trailing edges 140 , 142 and 144 , respectively that vary in distance from inlet 110 based on lateral position.
- Lobes 124 , 126 and 128 can vary in depth.
- inclination angle ⁇ 2 indicates the inclination of the bottom surface of a lobe with respect to second wall surface 104 (e.g., bottom surface 130 of lobe 124 in FIG. 4 ).
- Bottom surfaces 130 , 132 and 134 of respective lobes 124 , 126 and 128 can all have the same inclination angle ⁇ 2 and depth from second wall surface 104 .
- bottom surfaces 130 , 132 and 134 can have different inclination angles ⁇ 2 , forming lobes of differing depth.
- bottom surfaces 130 and 132 can have the same inclination angle ⁇ 2 while bottom surface 134 of middle lobe 128 has a different inclination angle ⁇ 2 and a depth different from lobes 124 and 126 .
- Lobes 124 , 126 and 128 can also vary in size. For example, as shown in FIG. 6 , lobes 126 and 128 are smaller (i.e. have smaller widths at an upstream region of diffusing section 114 ) than lobe 124 . In some embodiments, lobes 124 , 126 and 128 have the same size (e.g., same surface area). In alternate embodiments, lobes 124 , 126 and 128 have varying sizes or shapes to better laterally diffuse cooling air C according to the geometry of the component containing cooling hole 106 .
- Exemplary shapes and sizes of lobes 124 , 126 and 128 depend on a number of factors including: the thickness of wall 100 , the angle at which metering section 112 of cooling hole 106 is inclined relative to wall 100 , any curvature present on wall 100 in the vicinity of cooling hole 106 and/or the high temperature gas profile flowing past wall 100 .
- FIGS. 7 and 8 illustrate another embodiment of a cooling hole.
- FIG. 7 illustrates a sectional view of cooling hole 106 B (same cross section view as FIG. 4 ).
- FIG. 8 illustrates a view of cooling hole 106 B of FIG. 7 taken along the line 8 - 8 .
- diffusing section 114 also includes transition region 154 .
- ridge 146 (and lobes 124 and 126 ) do not extend all the way to outlet 116 . Instead, transition region 154 is positioned between outlet 116 and ridge 146 and lobes 124 and 126 .
- Transition region 154 can take various shapes and have different configurations depending on the location and desired flow profile of cooling hole 106 .
- the bottom surface of transition region 154 can be flat or curved.
- a curved (e.g., longitudinally convex) bottom surface of transition region 154 can facilitate improved flow attachment on the bottom surface.
- gas turbine engine components, gas path walls and cooling passages described herein can thus be manufactured using one or more of a variety of different processes. These techniques provide each cooling hole and cooling passage with its own particular configuration and features, including, but not limited to, inlet, metering, transition, diffusion, outlet, upstream wall, downstream wall, lateral wall, longitudinal, lobe and downstream edge features, as described above. In some cases, multiple techniques can be combined to improve overall cooling performance or reproducibility, or to reduce manufacturing costs.
- Suitable manufacturing techniques for forming the cooling configurations described here include, but are not limited to, electrical discharge machining (EDM), laser drilling, laser machining, electrical chemical machining (ECM), water jet machining, casting, conventional machining and combinations thereof.
- Electrical discharge machining includes both machining using a shaped electrode as well as multiple pass methods using a hollow spindle or similar electrode component.
- Laser machining methods include, but are not limited to, material removal by ablation, trepanning and percussion laser machining.
- Conventional machining methods include, but are not limited to, milling, drilling and grinding.
- the gas flow path walls and outer surfaces of some gas turbine engine components include one or more coatings, such as bond coats, thermal barrier coatings, abrasive coatings, abradable coatings and erosion or erosion-resistant coatings.
- coatings such as bond coats, thermal barrier coatings, abrasive coatings, abradable coatings and erosion or erosion-resistant coatings.
- the inlet, metering portion, transition, diffusion portion and outlet cooling features may be formed prior to coating application, after a first coating (e.g., a bond coat) is applied, or after a second or third (e.g., interlayer) coating process, or a final coating (e.g., environmental or thermal barrier) coating process.
- the diffusion portion and outlet features may be located within a wall or substrate, within a thermal barrier coating or other coating layer applied to a wall or substrate, or based on combinations thereof.
- the cooling geometry and other features may remain as described above, regardless of position relative to the wall and coating materials or airfoil materials.
- cooling features may affect selection of manufacturing techniques, including techniques used in forming the inlet, metering portion, transition, outlet, diffusion portion and other cooling features.
- a thermal barrier coat or other coating is applied to the outer surface of a gas path wall before the cooling hole or passage is produced.
- laser ablation or laser drilling may be used.
- either laser drilling or water jet machining may be used on a surface without a thermal barrier coat.
- different machining methods may be more or less suitable for forming different features of the cooling hole or cooling passage, for example, different EDM, laser machining and other machining techniques may be used for forming the outlet and diffusion features, and for forming the transition, metering and inlet features.
- FIG. 9A is a simplified flow diagram illustrating one embodiment of a method for producing a multi-lobed cooling hole in a gas turbine engine wall having first and second wall surfaces.
- Method 200 includes forming a metering section between the first and second surfaces (step 202 ) and forming a diffusing section between the metering section and the second wall surface (step 204 ).
- Metering section 112 is formed in step 202 by one or more of the casting, machining or drilling techniques described above. The technique(s) chosen is/are typically determined based on performance, reproducibility and cost.
- inlet 110 and portions of diffusing section 114 and outlet 116 can also be formed during formation of metering section 112 .
- Diffusing section 114 is formed in step 204 by one or more of the casting, machining or drilling techniques described above. As with metering section 112 , the technique(s) chosen is/are typically determined based on performance, reproducibility and cost.
- the diffusing section is formed in step 204 to have a first lobe in line with the metering section and a second lobe that diverges laterally from the metering section. Diffusing section 114 distributes the flow of the fluid into the lobes to form a film of cooling fluid at a hole outlet at the second wall surface of the gas turbine engine wall.
- step 202 occurs prior to step 204 , outlet 116 is fully formed once step 204 has been completed.
- Method 200 can be performed before or after an optional thermal barrier coating application.
- step 206 shown as a step in method 200 A in FIG. 9B
- a thermal barrier coating is applied to second wall surface 104 .
- Application of the thermal barrier coating can also include the application of a bond coating prior to the thermal barrier coating.
- the steps of method 200 A can be performed in any order depending on the location of cooling hole 106 and the location of diffusing section 114 relative to the metallic wall and the thermal barrier coating. As previously stated, the order of the steps can affect the machining or drilling techniques chosen.
- a gas turbine engine component can include a wall having first and second wall surfaces and a cooling hole extending through the wall.
- the cooling hole can include an inlet located at the first wall surface, an outlet located at the second wall surface, a metering section extending downstream from the inlet and a diffusing section extending from the metering section to the outlet.
- the diffusing section can include a first lobe diverging longitudinally from the metering section and a second lobe adjacent the first lobe and diverging longitudinally and laterally from the metering section.
- the system of the preceding paragraph can optionally include, additionally and/or alternatively any, one or more of the following features, configurations and/or additional components:
- At least one of the first and second lobes can include a curved bottom portion
- the first lobe and the second lobe can meet at a ridge
- At least one of the first and second lobes can include a curved outer portion
- the metering section can include a first lateral side, and the first lateral side of the metering section can be parallel to the curved outer portion of the first lobe;
- the metering section can further include a longitudinal axis, and the curved outer portion of the first lobe and the first lateral side of the metering section can be equidistant from the longitudinal axis of the metering section;
- the metering section can be inclined between the first wall surface and the second wall surface
- the first lobe can include a first depth and a first downstream angle
- the second lobe can include a second depth and a second downstream angle
- the first depth and the second depth can be equal and the first downstream angle and the second downstream angle can be equal
- the first lobe can include a first depth and a first downstream angle
- the second lobe can include a second depth and a second downstream angle
- the first depth and the second depth can be different or the first downstream angle and the second downstream angle can be different
- the diffusing section can further include a transition region extending between the first and second lobes and the outlet;
- the transition region can further include a curved surface;
- the component can be selected from the group consisting of blades, vanes, airfoil tips, airfoil platforms, combustors, blade outer air seals and augmentors.
- a wall of a component of a gas turbine engine can include first and second wall surfaces, an inlet located at the first wall surface, an outlet located at the second wall surface, a metering section commencing at the inlet and extending downstream from the inlet and a diffusing section extending from the metering section and terminating at the outlet.
- the diffusing section can include a first lobe diverging longitudinally from the metering section, a second lobe adjacent the first lobe and diverging longitudinally and laterally from the metering section and a ridge located between the first and second lobes.
- the system of the preceding paragraph can optionally include, additionally and/or alternatively any, one or more of the following features, configurations and/or additional components:
- At least one of the first and second lobes can include a curved bottom portion
- At least one of the first and second lobes can include a curved outer portion
- the metering section can include a first lateral side, and the first lateral side of the metering section can be parallel to the curved outer portion of the first lobe;
- the metering section can further include a longitudinal axis, and the curved outer portion of the first lobe and the first lateral side of the metering section can be equidistant from the longitudinal axis of the metering section;
- the diffusing section can further include a transition region extending between the first and second lobes and the outlet;
- the transition region can further include a curved surface;
- the component can be selected from the group consisting of blades, vanes, airfoil tips, airfoil platforms, combustors, blade outer air seals and augmentors.
- a method for producing a cooling hole in a gas turbine engine wall having first and second wall surfaces can include forming a metering section between the first wall surface and the second wall surface and forming a diffusing section between the metering section and the second wall surface.
- the diffusing section can include a first lobe in line with the metering section and a second lobe that diverges laterally from the metering section.
- the diffusing section distributes the flow of the fluid into the lobes to form a film of cooling fluid at a hole outlet at the second wall surface of the gas turbine engine wall.
- the system of the preceding paragraph can optionally include, additionally and/or alternatively any, one or more of the following features, configurations and/or additional components:
- forming the metering section and forming the diffusing section can be performed by electrical discharge machining, laser drilling, laser machining, electrical chemical machining, waterjet machining, casting, conventional machining and combinations thereof.
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Abstract
Description
- This application claims priority to U.S. Provisional Application No. 61/599,387, filed on Feb. 15, 2012 and entitled “COOLING HOLE WITH ASYMMETRIC DIFFUSER”, U.S. Provisional Application No. 61/599,381, filed on Feb. 15, 2012 and entitled “TRI-LOBED COOLING HOLE AND METHOD OF MANUFACTURE”, U.S. Provisional Application No. 61/599,372, filed on Feb. 15, 2012 and entitled “MULTI-LOBED COOLING HOLE AND METHOD OF MANUFACTURE”, the disclosures of which are incorporated by reference in their entirety.
- This invention relates generally to turbomachinery, and specifically to turbine flow path components for gas turbine engines. In particular, the invention relates to cooling techniques for airfoils and other gas turbine engine components exposed to hot working fluid flow, including, but not limited to, rotor blades and stator vane airfoils, endwall surfaces including platforms, shrouds and compressor and turbine casings, combustor liners, turbine exhaust assemblies, thrust augmentors and exhaust nozzles.
- Gas turbine engines are rotary-type combustion turbine engines built around a power core made up of a compressor, combustor and turbine, arranged in flow series with an upstream inlet and downstream exhaust. The compressor section compresses air from the inlet, which is mixed with fuel in the combustor and ignited to generate hot combustion gas. The turbine section extracts energy from the expanding combustion gas, and drives the compressor section via a common shaft. Expanded combustion products are exhausted downstream, and energy is delivered in the form of rotational energy in the shaft, reactive thrust from the exhaust, or both.
- Gas turbine engines provide efficient, reliable power for a wide range of applications in aviation, transportation and industrial power generation. Small-scale gas turbine engines typically utilize a one-spool design, with co-rotating compressor and turbine sections. Larger-scale combustion turbines including jet engines and industrial gas turbines (IGTs) are generally arranged into a number of coaxially nested spools. The spools operate at different pressures, temperatures and spool speeds, and may rotate in different directions.
- Individual compressor and turbine sections in each spool may also be subdivided into a number of stages, formed of alternating rows of rotor blade and stator vane airfoils. The airfoils are shaped to turn, accelerate and compress the working fluid flow, or to generate lift for conversion to rotational energy in the turbine.
- Industrial gas turbines often utilize complex nested spool configurations, and deliver power via an output shaft coupled to an electrical generator or other load, typically using an external gearbox. In combined cycle gas turbines (CCGTs), a steam turbine or other secondary system is used to extract additional energy from the exhaust, improving thermodynamic efficiency. Gas turbine engines are also used in marine and land-based applications, including naval vessels, trains and armored vehicles, and in smaller-scale applications such as auxiliary power units.
- Aviation applications include turbojet, turbofan, turboprop and turboshaft engine designs. In turbojet engines, thrust is generated primarily from the exhaust. Modern fixed-wing aircraft generally employ turbofan and turboprop configurations, in which the low pressure spool is coupled to a propulsion fan or propeller. Turboshaft engines are employed on rotary-wing aircraft, including helicopters, typically using a reduction gearbox to control blade speed. Unducted (open rotor) turbofans and ducted propeller engines also known, in a variety of single-rotor and contra-rotating designs with both forward and aft mounting configurations.
- Aviation turbines generally utilize two and three-spool configurations, with a corresponding number of coaxially rotating turbine and compressor sections. In two-spool designs, the high pressure turbine drives a high pressure compressor, forming the high pressure spool or high spool. The low-pressure turbine drives the low spool and fan section, or a shaft for a rotor or propeller. In three-spool engines, there is also an intermediate pressure spool. Aviation turbines are also used to power auxiliary devices including electrical generators, hydraulic pumps and elements of the environmental control system, for example using bleed air from the compressor or via an accessory gearbox.
- Additional turbine engine applications and turbine engine types include intercooled, regenerated or recuperated and variable cycle gas turbine engines, and combinations thereof. In particular, these applications include intercooled turbine engines, for example with a relatively higher pressure ratio, regenerated or recuperated gas turbine engines, for example with a relatively lower pressure ratio or for smaller-scale applications, and variable cycle gas turbine engines, for example for operation under a range of flight conditions including subsonic, transonic and supersonic speeds. Combined intercooled and regenerated/recuperated engines are also known, in a variety of spool configurations with traditional and variable cycle modes of operation.
- Turbofan engines are commonly divided into high and low bypass configurations. High bypass turbofans generate thrust primarily from the fan, which accelerates airflow through a bypass duct oriented around the engine core. This design is common on commercial aircraft and transports, where noise and fuel efficiency are primary concerns. The fan rotor may also operate as a first stage compressor, or as a pre-compressor stage for the low-pressure compressor or booster module. Variable-area nozzle surfaces can also be deployed to regulate the bypass pressure and improve fan performance, for example during takeoff and landing. Advanced turbofan engines may also utilize a geared fan drive mechanism to provide greater speed control, reducing noise and increasing engine efficiency, or to increase or decrease specific thrust.
- Low bypass turbofans produce proportionally more thrust from the exhaust flow, generating greater specific thrust for use in high-performance applications including supersonic jet aircraft. Low bypass turbofan engines may also include variable-area exhaust nozzles and afterburner or augmentor assemblies for flow regulation and short-term thrust enhancement. Specialized high-speed applications include continuously afterburning engines and hybrid turbojet/ramjet configurations.
- Across these applications, turbine performance depends on the balance between higher pressure ratios and core gas path temperatures, which tend to increase efficiency, and the related effects on service life and reliability due to increased stress and wear. This balance is particularly relevant to gas turbine engine components in the hot sections of the compressor, combustor, turbine and exhaust sections, where active cooling is required to prevent damage due to high gas path temperatures and pressures.
- Components present in the hot gas path of a gas turbine engine require cooling to prevent component melting and to reduce the effects of thermal fatigue and wear. Hollow blades and vanes, combustor walls and other components include thin metal walls made of high strength materials that provide durability. While these materials reduce the amount of cooling necessary, components in the hot gas path still require some sort of surface cooling.
- Film cooling holes are often used to cool these components. This type of cooling works by delivering cool air (e.g., air bled from a compressor) through small holes in the wall surface of the component. This air creates a thin layer (film) of cool air on the surface of the component wall, protecting it from higher temperature air and gases. One consideration with film cooling is that injecting cool air into a component reduces engine efficiency. The drop in efficiency increases as the amount of cooling airflow increases.
- Diffusion cooling holes were designed to increase the spread of the cooling film to reduce the debit on engine efficiency. By spreading out the film of cooling air, smaller amounts of cooling air could be used to cool an area. One problem with diffusion cooling holes is flow separation. Diffusion cooling holes can only spread cooling air to a certain extent before the flow separates, creating a “hole” in the cooling film. Flow separation is likely to occur at the “corners” of state of the art diffusion holes. Additionally, at high blowing ratios, the cooling film can “jet” or “blow off” the surface of the component, allowing nearby hot gases to cover the surface and reducing cooling effectiveness.
- A gas turbine engine component includes a wall having first and second wall surfaces and a cooling hole extending through the wall. The cooling hole includes an inlet located at the first wall surface, an outlet located at the second wall surface, a metering section extending downstream from the inlet and a diffusing section extending from the metering section to the outlet. The diffusing section includes a first lobe diverging longitudinally from the metering section and a second lobe adjacent the first lobe and diverging longitudinally and laterally from the metering section.
- A wall of a component of a gas turbine engine includes first and second wall surfaces, an inlet located at the first wall surface, an outlet located at the second wall surface, a metering section commencing at the inlet and extending downstream from the inlet and a diffusing section extending from the metering section and terminating at the outlet. The diffusing section includes a first lobe diverging longitudinally from the metering section, a second lobe adjacent the first lobe and diverging longitudinally and laterally from the metering section and a ridge located between the first and second lobes.
- A method for producing a cooling hole in a gas turbine engine wall having first and second wall surfaces includes forming a metering section between the first wall surface and the second wall surface and forming a diffusing section between the metering section and the second wall surface. The diffusing section includes a first lobe in line with the metering section and a second lobe that diverges laterally from the metering section. The diffusing section distributes the flow of the fluid into the lobes to form a film of cooling fluid at a hole outlet at the second wall surface of the gas turbine engine wall.
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FIG. 1 is a cross-sectional view of a gas turbine engine. -
FIG. 2A is a perspective view of an airfoil for the gas turbine engine, in a rotor blade configuration. -
FIG. 2B is a perspective view of an airfoil for the gas turbine engine, in a stator vane configuration. -
FIG. 3 is a view of a wall having cooling holes with asymmetric diffusing sections. -
FIG. 4 is a sectional view of the cooling hole ofFIG. 3 taken along the line 4-4. -
FIG. 5 is a view of the cooling hole ofFIG. 4 taken along the line 5-5. -
FIG. 5A is a view of the cooling hole ofFIG. 5 taken along the line A-A. -
FIG. 6 is another embodiment of a cooling hole with an asymmetric diffusing section. -
FIG. 7 is a sectional view of another embodiment of a cooling hole with an asymmetric diffusing section. -
FIG. 8 is a view of the cooling hole ofFIG. 7 taken along the line 8-8. -
FIG. 9A is a simplified flow diagram illustrating one embodiment of a method for producing a tri-lobed cooling hole. -
FIG. 9B is a simplified flow diagram illustrating another embodiment of a method for producing a tri-lobed cooling hole. -
FIG. 1 is a cross-sectional view ofgas turbine engine 10. Gas turbine engine (or turbine engine) 10 includes a power core withcompressor section 12,combustor 14 andturbine section 16 arranged in flow series betweenupstream inlet 18 anddownstream exhaust 20.Compressor section 12 andturbine section 16 are arranged into a number of alternating stages of rotor airfoils (or blades) 22 and stator airfoils (or vanes) 24. - In the turbofan configuration of
FIG. 1 ,propulsion fan 26 is positioned inbypass duct 28, which is coaxially oriented about the engine core along centerline (or turbine axis) CL. An open-rotor propulsion stage 26 may also provided, withturbine engine 10 operating as a turboprop or unducted turbofan engine. Alternatively,fan rotor 26 andbypass duct 28 may be absent, withturbine engine 10 configured as a turbojet or turboshaft engine, or an industrial gas turbine. - For improved service life and reliability, components of
gas turbine engine 10 are provided with an improved cooling configuration, as described below. Suitable components for the cooling configuration includerotor airfoils 22,stator airfoils 24 and other gas turbine engine components exposed to hot gas flow, including, but not limited to, platforms, shrouds, casings and other endwall surfaces in hot sections ofcompressor 12 andturbine 16, and liners, nozzles, afterburners, augmentors and other gas wall components incombustor 14 andexhaust section 20. - In the two-spool, high bypass configuration of
FIG. 1 ,compressor section 12 includes low pressure compressor (LPC) 30 and high pressure compressor (HPC) 32, andturbine section 16 includes high pressure turbine (HPT) 34 and low pressure turbine (LPT) 36.Low pressure compressor 30 is rotationally coupled tolow pressure turbine 36 via low pressure (LP)shaft 38, forming the LP spool or low spool.High pressure compressor 32 is rotationally coupled tohigh pressure turbine 34 via high pressure (HP)shaft 40, forming the HP spool or high spool. - Flow F at
inlet 18 divides into primary (core) flow FP and secondary (bypass) flow FS downstream offan rotor 26.Fan rotor 26 accelerates secondary flow FS throughbypass duct 28, with fan exit guide vanes (FEGVs) 42 to reduce swirl and improve thrust performance. In some designs, structural guide vanes (SGVs) 42 are used, providing combined flow turning and load bearing capabilities. - Primary flow FP is compressed in
low pressure compressor 30 andhigh pressure compressor 32, then mixed with fuel incombustor 14 and ignited to generate hot combustion gas. The combustion gas expands to provide rotational energy inhigh pressure turbine 34 andlow pressure turbine 36, drivinghigh pressure compressor 32 andlow pressure compressor 30, respectively. Expanded combustion gases exit through exhaust section (or exhaust nozzle) 20, which can be shaped or actuated to regulate the exhaust flow and improve thrust performance. -
Low pressure shaft 38 andhigh pressure shaft 40 are mounted coaxially about centerline CL, and rotate at different speeds. Fan rotor (or other propulsion stage) 26 is rotationally coupled tolow pressure shaft 38. In advanced designs, fandrive gear system 44 is provided for additional fan speed control, improving thrust performance and efficiency with reduced noise output. -
Fan rotor 26 may also function as a first-stage compressor forgas turbine engine 10, andLPC 30 may be configured as an intermediate compressor or booster. Alternatively,propulsion stage 26 has an open rotor design, or is absent, as described above.Gas turbine engine 10 thus encompasses a wide range of different shaft, spool and turbine engine configurations, including one, two and three-spool turboprop and (high or low bypass) turbofan engines, turboshaft engines, turbojet engines, and multi-spool industrial gas turbines. - In each of these applications, turbine efficiency and performance depend on the overall pressure ratio, defined by the total pressure at
inlet 18 as compared to the exit pressure ofcompressor section 12, for example at the outlet ofhigh pressure compressor 32, enteringcombustor 14. Higher pressure ratios, however, also result in greater gas path temperatures, increasing the cooling loads onrotor airfoils 22,stator airfoils 24 and other components ofgas turbine engine 10. To reduce operating temperatures, increase service life and maintain engine efficiency, these components are provided with improved cooling configurations, as described below. Suitable components include, but are not limited to, cooled gas turbine engine components incompressor sections combustor 14,turbine sections exhaust section 20 ofgas turbine engine 10. -
FIG. 2A is a perspective view of rotor airfoil (or blade) 22 forgas turbine engine 10, as shown inFIG. 1 , or for another turbomachine.Rotor airfoil 22 extends axially from leadingedge 51 to trailingedge 52, defining pressure surface 53 (front) and suction surface 54 (back) therebetween. - Pressure and suction surfaces 53 and 54 form the major opposing surfaces or walls of
airfoil 22, extending axially between leadingedge 51 and trailingedge 52, and radially fromroot section 55, adjacent inner diameter (ID)platform 56, to tipsection 57,opposite ID platform 56. In some designs,tip section 57 is shrouded. - Cooling holes or
outlets 60 are provided on one or more surfaces ofairfoil 22, for example along leadingedge 51, trailingedge 52, pressure (or concave)surface 53, or suction (or convex)surface 54, or a combination thereof. Cooling holes orpassages 60 may also be provided on the endwall surfaces ofairfoil 22, for example alongID platform 56, or on a shroud or engine casingadjacent tip section 57. -
FIG. 2B is a perspective view of stator airfoil (or vane) 24 forgas turbine engine 10, as shown inFIG. 1 , or for another turbomachine.Stator airfoil 24 extends axially from leadingedge 61 to trailingedge 62, defining pressure surface 63 (front) and suction surface 64 (back) therebetween. Pressure and suction surfaces 63 and 64 extend from inner (or root)section 65,adjacent ID platform 66, to outer (or tip)section 67, adjacent outer diameter (OD)platform 68. - Cooling holes or
outlets 60 are provided along one or more surfaces ofairfoil 24, for example leading or trailingedge surface passages 60 may also be provided on the endwall surfaces ofairfoil 24, for example alongID platform 66 andOD platform 68. - Rotor airfoils 22 (
FIG. 2A ) and stator airfoils 24 (FIG. 2B ) are formed of high strength, heat resistant materials such as high temperature alloys and superalloys, and are provided with thermal and erosion-resistant coatings.Airfoils cooling holes 60 to reduce thermal fatigue and wear, and to prevent melting when exposed to hot gas flow in the higher temperature regions of a gas turbine engine or other turbomachine. Cooling holes 60 deliver cooling fluid (e.g., steam or air from a compressor) through the outer walls and platform structures ofairfoils - While surface cooling extends service life and increases reliability, injecting cooling fluid into the gas path also reduces engine efficiency, and the cost in efficiency increases with the required cooling flow. Cooling holes 60 are thus provided with improved metering and inlet geometry to reduce jets and blow off, and improved diffusion and exit geometry to reduce flow separation and corner effects. Cooling holes 60 reduce flow requirements and improve the spread of cooling fluid across the hot outer surfaces of
airfoils -
FIG. 3 illustrates a view of a wall of a gas turbine engine component having cooling holes with asymmetric diffusing sections.Wall 100 includesfirst wall surface 102 andsecond wall surface 104. As described in greater detail below,wall 100 is primarily metallic andsecond wall surface 104 can include a thermal barrier coating. Coolingholes 106 are oriented so that their inlets are positioned on thefirst wall surface 102 and their outlets are positioned onsecond wall surface 104. During gas turbine engine operation,second wall surface 104 is in proximity to high temperature gases (e.g., combustion gases, hot air). Cooling air is delivered insidewall 100 where it exits the interior of the component throughcooling holes 106 and forms a cooling film onsecond wall surface 104. As shown inFIG. 3 , cooling holes 106 have two lobes in the diffusing section of the cooling hole outlet onsecond wall surface 104. - As described below in greater detail, cooling air flows out of
cooling holes 106 and flows through each of the lobes in the diffusing section. Coolingholes 106 can be arranged in a row onwall 100 as shown inFIG. 3 and positioned axially so that the cooling air flows in substantially the same direction longitudinally as the high temperature gases flowingpast wall 100. In this embodiment, cooling air passing throughcooling holes 106 exits cooling holes traveling in substantially the same direction as the high temperature gases flowing along second wall surface 104 (represented by arrow H). Here, the linear row of cooling holes 106 is substantially perpendicular to the direction of flow H. In alternate embodiments, the orientation ofcooling holes 106 can be arranged onsecond wall surface 104 so that the flow of cooling air is at an angle between parallel and perpendicular. Coolingholes 106 can also be provided in a staggered formation onwall 100. Coolingholes 106 can be located on a variety of components that require cooling. Suitable components include, but are not limited to, turbine vanes and blades, combustors, blade outer air seals, augmentors, etc. Coolingholes 106 can be located on the pressure side or suction side of vanes and blades. Coolingholes 106 can also be located on the blade tip or blade or vane platforms. Coolingholes 106 can also be located near airfoil endwalls or at other locations and individually aligned to provide targeted flow of cooling air. -
FIGS. 4 and 5 illustrate embodiments ofcooling hole 106 in greater detail.FIG. 4 illustrates a sectional view ofcooling hole 106 ofFIG. 3 taken along the line 4-4.FIG. 5 illustrates a view ofcooling hole 106 ofFIG. 4 taken along the line 5-5.Cooling hole 106 includesinlet 110,metering section 112 and diffusingsection 114.Inlet 110 is an opening located onfirst wall surface 102. Cooling air C enters coolinghole 106 throughinlet 110 and passes throughmetering section 112 and diffusingsection 114 before exitingcooling hole 106 atoutlet 116 alongsecond wall surface 104. -
Metering section 112 extends downstream frominlet 110 and controls (meters) the flow of cooling air throughcooling hole 106. In exemplary embodiments,metering section 112 has a substantially constant flow area frominlet 110 to diffusingsection 114.Metering section 112 can have circular, oblong (oval or elliptical) or racetrack (oval with two parallel sides having straight portions) shaped cross sections. InFIGS. 4 and 5 ,metering section 112 has a circular cross section.Circular metering sections 112 have a length l and diameter d. In some embodiments,circular metering section 112 has a length l according to the relationship: d≦l≦3d. That is, the length ofmetering section 112 is between one and three times its diameter. The length ofmetering section 22 can exceed 3d, reaching upwards of 30d. In alternate embodiments,metering section 112 has an oblong or racetrack-shaped or other shaped cross section. As oblong and racetrack configurations are not circular, theirmetering sections 112 have a length l and hydraulic diameter dh. In some embodiments,metering section 112 has a length l according to the relationship: dh≦l≦3dh. That is, the length ofmetering section 112 is between one and three times its hydraulic diameter. Again, the length ofmetering section 112 can exceed 3dh, reaching upwards of 30dh. In exemplary embodiments,metering section 112 is inclined with respect towall 100 as illustrated inFIG. 4 (i.e.metering section 112 is not perpendicular to wall 100).Metering section 112 has a longitudinal axis represented bynumeral 118.Metering section 112 also has alateral sidewall 113 as shown inFIG. 5 . -
Diffusing section 114 is adjacent to and downstream frommetering section 112. Cooling air C diffuses within diffusingsection 114 before exitingcooling hole 106 atoutlet 116 alongsecond wall surface 104. Once cooling air C exitsmetering section 112, the flow of air expands to fill diffusingsection 114. Cooling air C diffuses longitudinally (shown best inFIG. 4 ). In some embodiments, cooling air diffuses both longitudinally and laterally (shown best inFIG. 5 ) in diffusingsection 114.Second wall surface 104 includes upstream end 120 (upstream of cooling hole 106) and downstream end 122 (downstream from cooling hole 106).Diffusing section 114 opens alongsecond wall surface 104 betweenupstream end 120 anddownstream end 122. As shown inFIG. 4 , cooling air C diffuses in diffusingsection 114 as it flows towardsoutlet 116. - As shown best in
FIG. 5 , diffusingsection 114 includes twolobes lobe Lobes side walls lobe edges Lobes ridge 146.Ridge 146 can be straight or curved, both longitudinally and laterally. As shown inFIG. 4 , each lobe diverges longitudinally frommetering section 112.FIG. 4 illustrates a sectional view taken through the center ofcooling hole 106 and showsridge 146 betweenlobes Ridge 146 is inclined with respect tosecond wall surface 104 as shown by inclination angle θ1.Bottom surfaces lobes second wall surface 104 as shown by inclination angle θ2. Inclination angle θ2 indicates a downstream angle for each lobe. In the embodiment shown inFIG. 4 , bottom surfaces 130 and 132 oflobes section 114 diverges longitudinally fromlongitudinal axis 118 as it “attaches” tobottom surfaces respective lobes Lobes second wall surface 104 at trailingedges - In some embodiments, cooling air C passing through
cooling hole 106 also diffuses longitudinally nearupstream end 120. The upstream portion of diffusingsection 114 is bounded byforward edge 150. Forward edge 150 can be parallel with the upstream edge of metering section 112 (and with longitudinal axis 118), inclined towardsupstream end 120 or inclined towardsdownstream end 122. In exemplary embodiments,forward edge 150 is parallel with the upstream edge of metering section 112 (i.e. no upstream longitudinal diffusion) or inclined towardsdownstream end 122. In the embodiment illustrated inFIG. 4 ,forward edge 150 is inclined slightly towardsupstream end 120 from longitudinal axis 118 (represented by inclination angle θ3). In some embodiments,forward edge 150 is inclined towardsupstream end 120 to accommodate certain manufacturing methods. In these embodiments, the magnitude of inclination angle θ3 is minimized to less than about 15° and, in another embodiment, to less than about 1°. By minimizing inclination angle θ3 and positioning the end offorward edge 150 atsecond wall surface 104 as far downstream as possible, cooling airC exiting outlet 116 is likely to be more effective. In some embodiments,forward edge 150 is inclined towards downstream 122 (rather than upstream end 120) at an inclination angle θ3 of up to about −2°. - While cooling air C diffuses longitudinally within diffusing
section 114 as shown inFIG. 4 , cooling air also diffuses laterally within diffusingsection 114 as shown inFIG. 5 .Lobe 126 diverges laterally with respect tometering section 112.Lobe 124 includesside wall 136 on the side oflobe 124opposite ridge 146.Lobe 126 includesside wall 138 on the side oflobe 126opposite ridge 148. As illustrated inFIG. 5 ,lobe 126 diverges laterally in a downward direction away fromcenterline axis 152.Centerline axis 152 is a longitudinal axis passing through the center ofmetering section 112. On the other hand,sidewall 136 oflobe 124 is parallel withlateral sidewall 113 ofmetering section 112 and does not diverge in an upward direction away fromcenterline axis 152. Thus,lobe 124 does not laterally diverge away fromcenterline axis 152 to a substantial degree.Ridge 146 is angled downward (as shown inFIG. 5 ) slightly, allowing some lateral divergence of flow throughlobe 124.Ridge 146 can be angled downward to a greater degree to increase the lateral divergence of flow throughlobe 124 in one direction. Alternatively,ridge 146 can be parallel tosidewall 136 andmetering section 112. -
Ridge 146 aids in directing cooling air C intolobes Ridge 146 is generally an inverted V-shaped portion where the adjacent lobes meet.Ridge 146 can form a sharp edge between the lobes, where edges of adjacent lobes meet at a point. Alternatively,ridge 146 can be rounded or have other geometric shapes.Ridge 146 can form a straight line between adjacent lobes. Alternatively,ridge 146 can be laterally curved. As cooling air C exitsmetering section 112 and enters diffusingsection 114, coolingair 26encounters ridge 146.Ridge 146 can extend farther towardssecond wall surface 104 thanlobes FIG. 4 and evidenced by the difference in inclination angles θ1 (top of ridge) and θ2 (bottom surface of lobe). As a result,ridge 146 projects towardssecond wall surface 104 and serve to guide the flow of cooling air C intolobes Ridge 146 divides the flow of cooling air C betweenlobes lobe 126 to diverge laterally to correspond to the shape oflobe 126. - In exemplary embodiments, bottom surfaces 130 and 132 of
lobes FIG. 5A , the outer portion oflobes Lobe 124 includes a curved surface atside wall 136 and acurved bottom surface 130.Lobe 126 includes a curved surface atside wall 138 and acurved bottom surface 132. In this embodiment, bottom surfaces 130 and 132 are concave (i.e. curve towards first wall surface 102). -
FIG. 6 is a top view of another embodiment of a cooling hole, coolinghole 106A. As shown inFIG. 6 , diffusingsection 114 ofcooling hole 106A includes three lobes.Lobe 128 is located betweenlobes Lobe 128 includesbottom surface 134 and trailingedge 144. Ridge 147 separateslobe 124 andlobe 128, andridge 148 separateslobe 126 andlobe 128. Addinglobe 128 increases the amount of lateral divergence of cooling air C in diffusingsection 114.Ridges 147 and 148 divide the flow of cooling air C betweenlobes lobes -
Lobes second wall surface 104 at trailingedges Lobes second wall surface 104 in a number of ways. In one embodiment, each lobe blends withsecond wall surface 104 at the same axial distance frominlet 110, such that trailingedges FIG. 5 illustrates an embodiment in which the trailing edges of the lobes form a generally straight line. In another embodiment, trailingedges upstream end 120. In the embodiment illustrated inFIG. 6 ,lobes edges inlet 110 based on lateral position. -
Lobes bottom surface 130 oflobe 124 inFIG. 4 ). Bottom surfaces 130, 132 and 134 ofrespective lobes second wall surface 104. Alternatively, bottom surfaces 130, 132 and 134 can have different inclination angles θ2, forming lobes of differing depth. For example, bottom surfaces 130 and 132 can have the same inclination angle θ2 whilebottom surface 134 ofmiddle lobe 128 has a different inclination angle θ2 and a depth different fromlobes -
Lobes FIG. 6 ,lobes lobe 124. In some embodiments,lobes lobes cooling hole 106. Exemplary shapes and sizes oflobes wall 100, the angle at whichmetering section 112 ofcooling hole 106 is inclined relative to wall 100, any curvature present onwall 100 in the vicinity of coolinghole 106 and/or the high temperature gas profile flowingpast wall 100. -
FIGS. 7 and 8 illustrate another embodiment of a cooling hole.FIG. 7 illustrates a sectional view ofcooling hole 106B (same cross section view asFIG. 4 ).FIG. 8 illustrates a view ofcooling hole 106B ofFIG. 7 taken along the line 8-8. In this embodiment, diffusingsection 114 also includestransition region 154. As shown inFIG. 7 , ridge 146 (andlobes 124 and 126) do not extend all the way tooutlet 116. Instead,transition region 154 is positioned betweenoutlet 116 andridge 146 andlobes Transition region 154 can take various shapes and have different configurations depending on the location and desired flow profile ofcooling hole 106. The bottom surface oftransition region 154 can be flat or curved. A curved (e.g., longitudinally convex) bottom surface oftransition region 154 can facilitate improved flow attachment on the bottom surface. - The gas turbine engine components, gas path walls and cooling passages described herein can thus be manufactured using one or more of a variety of different processes. These techniques provide each cooling hole and cooling passage with its own particular configuration and features, including, but not limited to, inlet, metering, transition, diffusion, outlet, upstream wall, downstream wall, lateral wall, longitudinal, lobe and downstream edge features, as described above. In some cases, multiple techniques can be combined to improve overall cooling performance or reproducibility, or to reduce manufacturing costs.
- Suitable manufacturing techniques for forming the cooling configurations described here include, but are not limited to, electrical discharge machining (EDM), laser drilling, laser machining, electrical chemical machining (ECM), water jet machining, casting, conventional machining and combinations thereof. Electrical discharge machining includes both machining using a shaped electrode as well as multiple pass methods using a hollow spindle or similar electrode component. Laser machining methods include, but are not limited to, material removal by ablation, trepanning and percussion laser machining. Conventional machining methods include, but are not limited to, milling, drilling and grinding.
- The gas flow path walls and outer surfaces of some gas turbine engine components include one or more coatings, such as bond coats, thermal barrier coatings, abrasive coatings, abradable coatings and erosion or erosion-resistant coatings. For components having a coating, the inlet, metering portion, transition, diffusion portion and outlet cooling features may be formed prior to coating application, after a first coating (e.g., a bond coat) is applied, or after a second or third (e.g., interlayer) coating process, or a final coating (e.g., environmental or thermal barrier) coating process. Depending on component type, cooling hole or passage location, repair requirements and other considerations, the diffusion portion and outlet features may be located within a wall or substrate, within a thermal barrier coating or other coating layer applied to a wall or substrate, or based on combinations thereof. The cooling geometry and other features may remain as described above, regardless of position relative to the wall and coating materials or airfoil materials.
- In addition, the order in which cooling features are formed and coatings are applied may affect selection of manufacturing techniques, including techniques used in forming the inlet, metering portion, transition, outlet, diffusion portion and other cooling features. For example, when a thermal barrier coat or other coating is applied to the outer surface of a gas path wall before the cooling hole or passage is produced, laser ablation or laser drilling may be used. Alternatively, either laser drilling or water jet machining may be used on a surface without a thermal barrier coat. Additionally, different machining methods may be more or less suitable for forming different features of the cooling hole or cooling passage, for example, different EDM, laser machining and other machining techniques may be used for forming the outlet and diffusion features, and for forming the transition, metering and inlet features.
-
FIG. 9A is a simplified flow diagram illustrating one embodiment of a method for producing a multi-lobed cooling hole in a gas turbine engine wall having first and second wall surfaces.Method 200 includes forming a metering section between the first and second surfaces (step 202) and forming a diffusing section between the metering section and the second wall surface (step 204).Metering section 112 is formed instep 202 by one or more of the casting, machining or drilling techniques described above. The technique(s) chosen is/are typically determined based on performance, reproducibility and cost. In embodiments wherestep 202 occurs prior to step 204,inlet 110 and portions of diffusingsection 114 andoutlet 116 can also be formed during formation ofmetering section 112.Diffusing section 114 is formed instep 204 by one or more of the casting, machining or drilling techniques described above. As withmetering section 112, the technique(s) chosen is/are typically determined based on performance, reproducibility and cost. The diffusing section is formed instep 204 to have a first lobe in line with the metering section and a second lobe that diverges laterally from the metering section.Diffusing section 114 distributes the flow of the fluid into the lobes to form a film of cooling fluid at a hole outlet at the second wall surface of the gas turbine engine wall. - In embodiments where
step 202 occurs prior to step 204,outlet 116 is fully formed oncestep 204 has been completed.Method 200 can be performed before or after an optional thermal barrier coating application. In optional step 206 (shown as a step inmethod 200A inFIG. 9B ), a thermal barrier coating is applied tosecond wall surface 104. Application of the thermal barrier coating can also include the application of a bond coating prior to the thermal barrier coating. The steps ofmethod 200A can be performed in any order depending on the location of coolinghole 106 and the location of diffusingsection 114 relative to the metallic wall and the thermal barrier coating. As previously stated, the order of the steps can affect the machining or drilling techniques chosen. - While the invention has been described with reference to exemplary embodiments, it will be understood by those skilled in the art that various changes may be made and equivalents may be substituted for elements thereof without departing from the scope of the invention. In addition, many modifications may be made to adapt a particular situation or material to the teachings of the invention without departing from the essential scope thereof. Therefore, it is intended that the invention not be limited to the particular embodiments disclosed, but that the invention will include all embodiments falling within the scope of the appended claims.
- The following are non-exclusive descriptions of possible embodiments of the present invention.
- A gas turbine engine component can include a wall having first and second wall surfaces and a cooling hole extending through the wall. The cooling hole can include an inlet located at the first wall surface, an outlet located at the second wall surface, a metering section extending downstream from the inlet and a diffusing section extending from the metering section to the outlet. The diffusing section can include a first lobe diverging longitudinally from the metering section and a second lobe adjacent the first lobe and diverging longitudinally and laterally from the metering section.
- The system of the preceding paragraph can optionally include, additionally and/or alternatively any, one or more of the following features, configurations and/or additional components:
- at least one of the first and second lobes can include a curved bottom portion;
- the first lobe and the second lobe can meet at a ridge;
- at least one of the first and second lobes can include a curved outer portion;
- the metering section can include a first lateral side, and the first lateral side of the metering section can be parallel to the curved outer portion of the first lobe;
- the metering section can further include a longitudinal axis, and the curved outer portion of the first lobe and the first lateral side of the metering section can be equidistant from the longitudinal axis of the metering section;
- the metering section can be inclined between the first wall surface and the second wall surface;
- the first lobe can include a first depth and a first downstream angle, the second lobe can include a second depth and a second downstream angle, and the first depth and the second depth can be equal and the first downstream angle and the second downstream angle can be equal;
- the first lobe can include a first depth and a first downstream angle, the second lobe can include a second depth and a second downstream angle, and the first depth and the second depth can be different or the first downstream angle and the second downstream angle can be different;
- the diffusing section can further include a transition region extending between the first and second lobes and the outlet;
- the transition region can further include a curved surface; and/or
- the component can be selected from the group consisting of blades, vanes, airfoil tips, airfoil platforms, combustors, blade outer air seals and augmentors.
- A wall of a component of a gas turbine engine can include first and second wall surfaces, an inlet located at the first wall surface, an outlet located at the second wall surface, a metering section commencing at the inlet and extending downstream from the inlet and a diffusing section extending from the metering section and terminating at the outlet. The diffusing section can include a first lobe diverging longitudinally from the metering section, a second lobe adjacent the first lobe and diverging longitudinally and laterally from the metering section and a ridge located between the first and second lobes.
- The system of the preceding paragraph can optionally include, additionally and/or alternatively any, one or more of the following features, configurations and/or additional components:
- at least one of the first and second lobes can include a curved bottom portion;
- at least one of the first and second lobes can include a curved outer portion;
- the metering section can include a first lateral side, and the first lateral side of the metering section can be parallel to the curved outer portion of the first lobe;
- the metering section can further include a longitudinal axis, and the curved outer portion of the first lobe and the first lateral side of the metering section can be equidistant from the longitudinal axis of the metering section;
- the diffusing section can further include a transition region extending between the first and second lobes and the outlet;
- the transition region can further include a curved surface; and/or
- the component can be selected from the group consisting of blades, vanes, airfoil tips, airfoil platforms, combustors, blade outer air seals and augmentors.
- A method for producing a cooling hole in a gas turbine engine wall having first and second wall surfaces can include forming a metering section between the first wall surface and the second wall surface and forming a diffusing section between the metering section and the second wall surface. The diffusing section can include a first lobe in line with the metering section and a second lobe that diverges laterally from the metering section. The diffusing section distributes the flow of the fluid into the lobes to form a film of cooling fluid at a hole outlet at the second wall surface of the gas turbine engine wall.
- The system of the preceding paragraph can optionally include, additionally and/or alternatively any, one or more of the following features, configurations and/or additional components:
- forming the metering section and forming the diffusing section can be performed by electrical discharge machining, laser drilling, laser machining, electrical chemical machining, waterjet machining, casting, conventional machining and combinations thereof.
Claims (22)
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EP13784812.3A EP2815107B1 (en) | 2012-02-15 | 2013-02-12 | Gas turbine engine component and corresponding method for producing a cooling hole |
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Cited By (22)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US20150027127A1 (en) * | 2013-07-24 | 2015-01-29 | Rolls-Royce Deutschland Ltd & Co Kg | Combustion chamber tile of a gas turbine |
US9145773B2 (en) | 2012-05-09 | 2015-09-29 | General Electric Company | Asymmetrically shaped trailing edge cooling holes |
US9175569B2 (en) | 2012-03-30 | 2015-11-03 | General Electric Company | Turbine airfoil trailing edge cooling slots |
CN105909317A (en) * | 2015-02-24 | 2016-08-31 | 通用电气公司 | Engine component |
US20170003026A1 (en) * | 2015-06-30 | 2017-01-05 | Rolls-Royce Corporation | Combustor tile |
US9945233B2 (en) | 2013-05-22 | 2018-04-17 | Kawasaki Jukogyo Kabushiki Kaisha | Double-jet film cooling structure and method for manufacturing same |
US9976746B2 (en) | 2015-09-02 | 2018-05-22 | General Electric Company | Combustor assembly for a turbine engine |
US10024169B2 (en) | 2015-02-27 | 2018-07-17 | General Electric Company | Engine component |
US10132166B2 (en) | 2015-02-27 | 2018-11-20 | General Electric Company | Engine component |
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US10168051B2 (en) | 2015-09-02 | 2019-01-01 | General Electric Company | Combustor assembly for a turbine engine |
US10197278B2 (en) | 2015-09-02 | 2019-02-05 | General Electric Company | Combustor assembly for a turbine engine |
US20190071976A1 (en) * | 2017-09-07 | 2019-03-07 | General Electric Company | Component for a turbine engine with a cooling hole |
US10386069B2 (en) | 2012-06-13 | 2019-08-20 | General Electric Company | Gas turbine engine wall |
US10392947B2 (en) | 2015-07-13 | 2019-08-27 | General Electric Company | Compositions and methods of attachment of thick environmental barrier coatings on CMC components |
US10494928B2 (en) | 2014-10-06 | 2019-12-03 | Rolls-Royce Plc | Cooled component |
US10563867B2 (en) | 2015-09-30 | 2020-02-18 | General Electric Company | CMC articles having small complex features for advanced film cooling |
US11149646B2 (en) | 2015-09-02 | 2021-10-19 | General Electric Company | Piston ring assembly for a turbine engine |
US11286791B2 (en) * | 2016-05-19 | 2022-03-29 | Honeywell International Inc. | Engine components with cooling holes having tailored metering and diffuser portions |
US11402097B2 (en) | 2018-01-03 | 2022-08-02 | General Electric Company | Combustor assembly for a turbine engine |
US11407065B2 (en) * | 2017-06-16 | 2022-08-09 | Raytheon Corporation Inc. | Systems and methods for manufacturing film cooling hole diffuser portion |
CN116085117A (en) * | 2023-04-10 | 2023-05-09 | 清华大学 | Guiding structure |
Families Citing this family (11)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US9024226B2 (en) * | 2012-02-15 | 2015-05-05 | United Technologies Corporation | EDM method for multi-lobed cooling hole |
US9598979B2 (en) | 2012-02-15 | 2017-03-21 | United Technologies Corporation | Manufacturing methods for multi-lobed cooling holes |
EP2956633B1 (en) * | 2013-02-15 | 2021-05-05 | Raytheon Technologies Corporation | Component for a gas turbine engine and corresponding method of forming a cooling hole |
EP2971669B1 (en) * | 2013-03-15 | 2018-12-19 | United Technologies Corporation | Wall of a component of a gas turbine engine and corresponding gas turbine engine component |
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US20160090843A1 (en) * | 2014-09-30 | 2016-03-31 | General Electric Company | Turbine components with stepped apertures |
US20160169004A1 (en) | 2014-12-15 | 2016-06-16 | United Technologies Corporation | Cooling passages for gas turbine engine component |
US10400607B2 (en) | 2014-12-30 | 2019-09-03 | United Technologies Corporation | Large-footprint turbine cooling hole |
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US11585224B2 (en) | 2020-08-07 | 2023-02-21 | General Electric Company | Gas turbine engines and methods associated therewith |
Citations (3)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US8057181B1 (en) * | 2008-11-07 | 2011-11-15 | Florida Turbine Technologies, Inc. | Multiple expansion film cooling hole for turbine airfoil |
US20120167389A1 (en) * | 2011-01-04 | 2012-07-05 | General Electric Company | Method for providing a film cooled article |
US8245519B1 (en) * | 2008-11-25 | 2012-08-21 | Florida Turbine Technologies, Inc. | Laser shaped film cooling hole |
Family Cites Families (62)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US4197443A (en) | 1977-09-19 | 1980-04-08 | General Electric Company | Method and apparatus for forming diffused cooling holes in an airfoil |
US4529358A (en) | 1984-02-15 | 1985-07-16 | The United States Of America As Represented By The Administrator Of The National Aeronautics And Space Administration | Vortex generating flow passage design for increased film cooling effectiveness |
US4653279A (en) | 1985-01-07 | 1987-03-31 | United Technologies Corporation | Integral refilmer lip for floatwall panels |
US4622821A (en) | 1985-01-07 | 1986-11-18 | United Technologies Corporation | Combustion liner for a gas turbine engine |
US4700544A (en) | 1985-01-07 | 1987-10-20 | United Technologies Corporation | Combustors |
US4653983A (en) | 1985-12-23 | 1987-03-31 | United Technologies Corporation | Cross-flow film cooling passages |
US4672727A (en) | 1985-12-23 | 1987-06-16 | United Technologies Corporation | Method of fabricating film cooling slot in a hollow airfoil |
US4738588A (en) | 1985-12-23 | 1988-04-19 | Field Robert E | Film cooling passages with step diffuser |
US4684323A (en) | 1985-12-23 | 1987-08-04 | United Technologies Corporation | Film cooling passages with curved corners |
US6139258A (en) | 1987-03-30 | 2000-10-31 | United Technologies Corporation | Airfoils with leading edge pockets for reduced heat transfer |
GB2227965B (en) | 1988-10-12 | 1993-02-10 | Rolls Royce Plc | Apparatus for drilling a shaped hole in a workpiece |
GB8830152D0 (en) | 1988-12-23 | 1989-09-20 | Rolls Royce Plc | Cooled turbomachinery components |
US5129231A (en) | 1990-03-12 | 1992-07-14 | United Technologies Corporation | Cooled combustor dome heatshield |
US5326224A (en) | 1991-03-01 | 1994-07-05 | General Electric Company | Cooling hole arrangements in jet engine components exposed to hot gas flow |
US6744010B1 (en) | 1991-08-22 | 2004-06-01 | United Technologies Corporation | Laser drilled holes for film cooling |
US5660525A (en) | 1992-10-29 | 1997-08-26 | General Electric Company | Film cooled slotted wall |
US5651662A (en) | 1992-10-29 | 1997-07-29 | General Electric Company | Film cooled wall |
US5252026A (en) | 1993-01-12 | 1993-10-12 | General Electric Company | Gas turbine engine nozzle |
US5419681A (en) | 1993-01-25 | 1995-05-30 | General Electric Company | Film cooled wall |
US5683600A (en) | 1993-03-17 | 1997-11-04 | General Electric Company | Gas turbine engine component with compound cooling holes and method for making the same |
US5358374A (en) | 1993-07-21 | 1994-10-25 | General Electric Company | Turbine nozzle backflow inhibitor |
US5382133A (en) | 1993-10-15 | 1995-01-17 | United Technologies Corporation | High coverage shaped diffuser film hole for thin walls |
US5418345A (en) | 1994-02-28 | 1995-05-23 | United Technologies Corporation | Method for forming shaped passages |
US5609779A (en) | 1996-05-15 | 1997-03-11 | General Electric Company | Laser drilling of non-circular apertures |
US6092982A (en) | 1996-05-28 | 2000-07-25 | Kabushiki Kaisha Toshiba | Cooling system for a main body used in a gas stream |
US5813836A (en) | 1996-12-24 | 1998-09-29 | General Electric Company | Turbine blade |
US6287075B1 (en) | 1997-10-22 | 2001-09-11 | General Electric Company | Spanwise fan diffusion hole airfoil |
EP0950463B1 (en) | 1998-03-23 | 2002-01-23 | Alstom | Non-circular cooling hole and method of manufacturing the same |
EP0945593B1 (en) | 1998-03-23 | 2003-05-07 | ALSTOM (Switzerland) Ltd | Film-cooling hole |
GB9821639D0 (en) | 1998-10-06 | 1998-11-25 | Rolls Royce Plc | Coolant passages for gas turbine components |
US6243948B1 (en) | 1999-11-18 | 2001-06-12 | General Electric Company | Modification and repair of film cooling holes in gas turbine engine components |
GB0001399D0 (en) | 2000-01-22 | 2000-03-08 | Rolls Royce Plc | An aerofoil for an axial flow turbomachine |
US6973419B1 (en) | 2000-03-02 | 2005-12-06 | United Technologies Corporation | Method and system for designing an impingement film floatwall panel system |
JP3782637B2 (en) | 2000-03-08 | 2006-06-07 | 三菱重工業株式会社 | Gas turbine cooling vane |
US6944580B1 (en) | 2000-06-30 | 2005-09-13 | United Technologies Corporation | Method and system for designing frames and cases |
US6561758B2 (en) | 2001-04-27 | 2003-05-13 | General Electric Company | Methods and systems for cooling gas turbine engine airfoils |
US6547524B2 (en) | 2001-05-21 | 2003-04-15 | United Technologies Corporation | Film cooled article with improved temperature tolerance |
US6494678B1 (en) | 2001-05-31 | 2002-12-17 | General Electric Company | Film cooled blade tip |
US6984102B2 (en) | 2003-11-19 | 2006-01-10 | General Electric Company | Hot gas path component with mesh and turbulated cooling |
JP3997986B2 (en) | 2003-12-19 | 2007-10-24 | 株式会社Ihi | Cooling turbine component and cooling turbine blade |
US7328580B2 (en) | 2004-06-23 | 2008-02-12 | General Electric Company | Chevron film cooled wall |
GB0424593D0 (en) | 2004-11-06 | 2004-12-08 | Rolls Royce Plc | A component having a film cooling arrangement |
US7186085B2 (en) | 2004-11-18 | 2007-03-06 | General Electric Company | Multiform film cooling holes |
US7883320B2 (en) | 2005-01-24 | 2011-02-08 | United Technologies Corporation | Article having diffuser holes and method of making same |
US7374401B2 (en) | 2005-03-01 | 2008-05-20 | General Electric Company | Bell-shaped fan cooling holes for turbine airfoil |
US20080003096A1 (en) | 2006-06-29 | 2008-01-03 | United Technologies Corporation | High coverage cooling hole shape |
US7887294B1 (en) | 2006-10-13 | 2011-02-15 | Florida Turbine Technologies, Inc. | Turbine airfoil with continuous curved diffusion film holes |
US7726131B2 (en) | 2006-12-19 | 2010-06-01 | Pratt & Whitney Canada Corp. | Floatwall dilution hole cooling |
US20080145208A1 (en) | 2006-12-19 | 2008-06-19 | General Electric Company | Bullnose seal turbine stage |
US7578653B2 (en) | 2006-12-19 | 2009-08-25 | General Electric Company | Ovate band turbine stage |
US7766609B1 (en) | 2007-05-24 | 2010-08-03 | Florida Turbine Technologies, Inc. | Turbine vane endwall with float wall heat shield |
US8800293B2 (en) | 2007-07-10 | 2014-08-12 | United Technologies Corporation | Floatwell panel assemblies and related systems |
US8128366B2 (en) | 2008-06-06 | 2012-03-06 | United Technologies Corporation | Counter-vortex film cooling hole design |
US8092177B2 (en) | 2008-09-16 | 2012-01-10 | Siemens Energy, Inc. | Turbine airfoil cooling system with diffusion film cooling hole having flow restriction rib |
US8328517B2 (en) | 2008-09-16 | 2012-12-11 | Siemens Energy, Inc. | Turbine airfoil cooling system with diffusion film cooling hole |
US7997868B1 (en) | 2008-11-18 | 2011-08-16 | Florida Turbine Technologies, Inc. | Film cooling hole for turbine airfoil |
US8038399B1 (en) | 2008-11-22 | 2011-10-18 | Florida Turbine Technologies, Inc. | Turbine rim cavity sealing |
US8319146B2 (en) | 2009-05-05 | 2012-11-27 | General Electric Company | Method and apparatus for laser cutting a trench |
US20110097191A1 (en) | 2009-10-28 | 2011-04-28 | General Electric Company | Method and structure for cooling airfoil surfaces using asymmetric chevron film holes |
US8857055B2 (en) | 2010-01-29 | 2014-10-14 | General Electric Company | Process and system for forming shaped air holes |
US8905713B2 (en) | 2010-05-28 | 2014-12-09 | General Electric Company | Articles which include chevron film cooling holes, and related processes |
US8672613B2 (en) | 2010-08-31 | 2014-03-18 | General Electric Company | Components with conformal curved film holes and methods of manufacture |
-
2012
- 2012-07-09 US US13/544,136 patent/US8733111B2/en active Active
-
2013
- 2013-02-12 WO PCT/US2013/025705 patent/WO2013165507A2/en active Application Filing
- 2013-02-12 EP EP13784812.3A patent/EP2815107B1/en active Active
Patent Citations (3)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US8057181B1 (en) * | 2008-11-07 | 2011-11-15 | Florida Turbine Technologies, Inc. | Multiple expansion film cooling hole for turbine airfoil |
US8245519B1 (en) * | 2008-11-25 | 2012-08-21 | Florida Turbine Technologies, Inc. | Laser shaped film cooling hole |
US20120167389A1 (en) * | 2011-01-04 | 2012-07-05 | General Electric Company | Method for providing a film cooled article |
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US9945233B2 (en) | 2013-05-22 | 2018-04-17 | Kawasaki Jukogyo Kabushiki Kaisha | Double-jet film cooling structure and method for manufacturing same |
US20150027127A1 (en) * | 2013-07-24 | 2015-01-29 | Rolls-Royce Deutschland Ltd & Co Kg | Combustion chamber tile of a gas turbine |
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US10161266B2 (en) | 2015-09-23 | 2018-12-25 | General Electric Company | Nozzle and nozzle assembly for gas turbine engine |
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US11286791B2 (en) * | 2016-05-19 | 2022-03-29 | Honeywell International Inc. | Engine components with cooling holes having tailored metering and diffuser portions |
US11407065B2 (en) * | 2017-06-16 | 2022-08-09 | Raytheon Corporation Inc. | Systems and methods for manufacturing film cooling hole diffuser portion |
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Also Published As
Publication number | Publication date |
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EP2815107A4 (en) | 2015-12-16 |
EP2815107B1 (en) | 2020-11-18 |
WO2013165507A3 (en) | 2014-01-16 |
WO2013165507A2 (en) | 2013-11-07 |
US8733111B2 (en) | 2014-05-27 |
EP2815107A2 (en) | 2014-12-24 |
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